The invention relates to aircraft gas turbine engine systems for supplying aircraft system bleed air and starting the aircraft engine. In particular, the invention relates to an aircraft gas turbine engine integrated system for efficiently performing these functions and also reducing drag on engine nacelles or other surfaces by using laminar flow boundary layer air as the same air supplied to the aircraft bleed air system.
Modern day aircraft use gas turbine engines which in addition to propulsion, provide secondary functions required by the aircraft systems. These secondary functions include electrical power, hydraulic power and aircraft bleed air. Bleed air is normally taken from the engine compressor and, after precooling with engine fan air in a heat exchanger, is delivered to various aircraft systems such as cowl and wing anti-ice system and the cabin pressurization and environmental control system for controlling cabin air freshness and temperature These two systems are generally referred to as the anti-ice and ECS systems respectively. As part of the aircraft air ducting system, air is also routed in reverse flow to the engine where it powers an air turbine engine starter. Air for engine starting can be obtained from a ground cart, an on-board auxiliary power unit or bleed air from another engine.
Extraction of aircraft bleed air from the engine compressor has adverse affects on the propulsion cycle and engine life. Air taken into the engine compressor incurs a ram drag penalty (loss of momentum). Engine net thrust is equal to engine exhaust momentum minus inlet ram drag. Engine turbine power is needed to compress air and account for compressor inefficiency. Therefore, extra fuel consumption is always associated with bleed air (air which does not produce thrust). This extra fuel burned in the engine combustor results in higher gas temperature delivered to the engine turbine and reduction of turbine blade life. Such penalties must be incurred in order that the engine turbine provide extra power associated with bleed air. It is not possible, without undue complexity, to always bleed the engine compressor stage which provides exactly the correct pressure needed for the aircraft anti-ice and ECS systems. Typically only two bleed ports are provided. Therefore, the result is to bleed air which exceeds minimum pressure requirements resulting in even higher penalty to the engine cycle than would be required by the aircraft systems. Most often the bleed air is not only at a higher than required pressure, it is also too hot. For reasons of fire safety, maximum bleed air temperature is usually limited to 450.degree. to 500.degree. F. Temperature control requires cooling the bleed air with a precooler. Most modern engines use fan air to cool compressor bleed air. Use of fan air imposes an additional penalty on fuel consumption. Further, the precooler is usually large and requires a fan air scoop which produces drag. A typical large turbofan engine will consume about 3% extra fuel and run at about 50.degree. F. hotter turbine temperature in order to provide aircraft system bleed air. The present invention addresses these problems and deficiencies characteristic of the prior art and conventional apparatus used to supply aircraft bleed air.
A second aspect of this invention concerns the engine air driven starter. Air starters are conventionally air powered turbines mounted to the engine accessory gearbox. The starter turbine rotates at very high speed and drives the engine through a planetary gear system during engine acceleration to just below idle speed. Once the engine lights it begins to develop its own power and, at a speed below idle, accelerates away from the starter. An overrunning mechanical clutch allows the starter to disengage and then the starter air is shutoff and the starter turbine comes to rest. During the remainder of the flight the starter is not used for any purpose and simply represents extra weight carried around by the aircraft. Within a very narrow flight profile of the aircraft, the starter can sometimes be used for emergency engine relight, but only at conditions where the windmill speed of the engine is low enough that the starter clutch can be engaged without damage due to what is referred to as crash engagement. Engine starters can not be used during normal aircraft cruise conditions; where the only means for relight is from the freely windmilling engine. One advantage of the present invention is that it permits operation of the air starter during all aircraft flight conditions thereby avoiding the delay in engine relight which can be associated with flight conditions unfavorable for fast windmill relights. The present invention further enhances the solution to the relight problem by using the starter turbine during all operation conditions as a means for improving the performance of the auxiliary bleed air compressor.
A third aspect of this invention concerns cooling the engine compartment. Cooling air is conventionally removed from the engine fan duct and used as ventilation cooling air for engine accessories mounted outside the main engine casing. This is particularly necessary for electronic controls and electrical components. Cooling air is also used in conjunction with compressor and turbine clearance control systems; systems which control the gap between rotating blade and adjacent casing walls to prevent rubs and excessive clearance. Another fuel saving advantage of one embodiment of the present invention is to use the turbine associated with the engine starter and auxiliary bleed air compressor as a means for cooling air which can then be used for engine compartment cooling, electronic control cooling or clearance control.
A fourth aspect of this invention relates to aerodynamic drag associated with engine nacelles, pylons and other aero-flowpath surfaces. As air flows on to and over a surface such as an engine nacelle it progressively builds up a low velocity boundary layer of increasing thickness. Within this boundary layer a portion of the velocity component of free stream total pressure is converted to increased static pressure. As the result of rise in static pressure, boundary layer thickness, and diffusion a point is reached where back pressure causes an otherwise laminar boundary layer to become turbulent. In the turbulent region a considerable amount of total pressure is converted to static temperature represented thermodynamically as an increase in entropy. By the time the boundary layer leaves the surface, or in the particular case of an aircraft gas turbine engine the end of the nacelle, an unrecoverable loss in total pressure has occurred. The large entropy rise associated with turbulence is at the expense of air momentum. Turbulence also gives rise to increased static pressure which may increase the intensity of rearward acting pressure force on the surface. Now if the boundary layer thickness is kept small, separation and turbulence will not occur and drag can be substantially reduced. One way to avoid increase in boundary thickness is to pump or bleed off boundary layer air through holes in the surface. Boundary layer pumps or compressors would be desirable from an aerodynamic standpoint but because of the relatively large air flow rates associated with effective boundary layer pumping or bleeding, the concept has not been adapted to modern aircraft and engines. Therefore, one problem this invention is directed at is to effectively and economically use the engine auxiliary compressor to pump and compress the laminar flow boundary layer air.
In order to maximize performance of the present invention, the preferred embodiment of the present invention addresses a problem relating to matching the auxiliary compressor's operating line (pressure ratio and flow) with the aircraft systems required schedule for pressure and flow, conventionally supplied by engine bleed air. Operating conditions always occur where the system wants higher pressures at lower flows then the compressor can deliver without stall (too low a flow for the operating speed needed to produce the required pressure). The compressor must be allowed to pass extra flow and avoid stall. However, this extra flow is an added loss in terms of both extra power into the compressor and added ram drag associated with bringing the air on board the engine (also extra fan power if the air source is the fan duct). The preferred embodiment of the present invention provides an economical fuel saving solution to this problem as opposed to just dumping the unused flow. The invention provides a means to extract power from this extra flow after its compression. Furthermore, after the extra flow passes through the air turbine which is also used to start the engine and after it is expanded and useful power has been extracted to help drive the compressor, the extra flow is used to cool the engine compartment, electronic control, or as part of the clearance control system.
It is, therefore, an object of the present invention to provide a more efficient and longer life aircraft gas turbine engine by reduction or elimination of engine compressor bleed and its associated fan air precooler.
Another object of the present invention is to provide a fuel efficient system for supplying compressed air to the aircraft anti-ice and ECS systems.
Another object of the present invention is to provide the engine with a quick and reliable in flight restart or relight capability.
Yet another object of the present invention is to provide the engine with a starter that avoids the need for crash engagement for in flight relight.
Another object of the present invention is to provide cooling air to cool the engine compartment, electronic control or as part of the clearance control system.
A further object of the present invention is to reduce aircraft drag in a fuel efficient manner.
A further object of the present invention is to reduce the cost and complexity of the aircraft gas turbine engine.
Yet a further object of the present invention is to simplify the valving and ducting associated with an aircraft gas turbine engine.
These objects and other features and advantages will become more readily apparent in the following description when taken in conjunction with the appended drawings.